Light structural elements of the kind also used in particular in the aerospace industry often consist of an outer skin reinforced on the inside by a two-dimensional bracing. For example, an aircraft fuselage with an outer skin reinforced by stringers and frames is assembled in this way. Special emphasis is placed on reducing the weight during the design of such light structural elements, however the light structural elements shall satisfy varying strength, fatigue and damage tolerance requirements depending on the respective application. In particular in aircraft construction, there are specific criteria relative to the damage tolerance characteristics of the light structural elements used in this area. At the same time, requirements for a light and cost-effective structural design must be taken into account.
Various requirements relating to different strength, fatigue and damage shall be satisfied during the design of light structural elements as a function of a special application. In particular in aircraft construction, particular emphases is placed on the damage tolerance of light structural elements, since this is important for the safety of passenger transport.
Aircraft fuselages are usually made out of interconnected skin panels which have reinforcing elements. Various known methods join individual skin panels, as well as for joining panels with reinforcing elements. Riveting, bonding and welding procedures are used for joining individual skin panels, as well as for joining skin panels with reinforcing elements.
Advantages offered by welding procedures over bonding have to do with the process automatability, higher quality and reproducibility of individual bonds. Structural elements welded in aircraft construction are resistant to corrosion and easy to repair. They have very good static properties, and allow for a reduction in weight. In addition, the structures fabricated through welding are mostly comparatively cost-effective to manufacture.
DE 196 39 667 and DE 198 44 035 disclose a structural element with a welded skin stringer design. Large-sized skin panels are used to weld on profiles (stringers and frames) via laser welding.
Such structural elements, which may be used in an aircraft, must also satisfy requirements placed on a requisite damage tolerance. In particular in the area of bonds between the skin and reinforcing profiles, integral structures (welded or milled) are highly sensitive. Given a cracking of the skin field, the crack usually propagates under the reinforcing profile without damaging the same, in the case of riveted or adhesively bonded reinforcing profiles.
Given an integral design (such as a welded or milled brace), there is a danger that the crack will also propagate in the reinforcing element. This effect reduces the life of the reinforcing elements, and diminishes the damage tolerant properties of integral structures. Damage tolerance may be increased by raising the skin thickness, using additional local reinforcing means, or by adjusting the skin thickness to load requirements. However, these measures increase the weight, and are cost-intensive.
DE 199 24 909 discloses a thickening in the foot area of the brace in an integral structural element for crack delay or deflection. However, since the thickening is integral, i.e., the brace and thickening are designed as a combined component, there is also the danger of continuous cracking over the entire area of the brace.
Another method for improving damage tolerance involves using additional reinforcements inside the reinforcing profiles (see DE 101 63 848) or on the bracing profiles (see DE 100 31 510) to avoid crack propagation (as in FIG. 1). Given crack propagation in the bracing element, the delay in crack advancement is minimal owing to a quick initiation of the crack inside the reinforcing element. In addition, materials with better mechanical properties are used for the reinforcing elements. On the other hand, these materials have a lower resistance to crack propagation. As a result, the reinforcing elements have a short lifespan after crack initiation.
A structural element 100 for an aircraft described in the prior reference is depicted with reference to FIG. 1.
The structural element 100 comprises a skin panel 101 and stringer element 102 (as the reinforcing element), which are integrally joined, i.e., consist of a single material. To improve the damage tolerance of the structural element 100, a space that incorporates a reinforcing element 103 is provided in the stringer 102. The object of the reinforcing element 103 is to prevent the crack from propagating once a crack has arisen.
However, rapid crack initiation 104 may take place given crack propagation 105 in the structural element 100, despite the provision of a reinforcing element 103. The rapid crack initiation 104 results in a break in the reinforcing element 103. As a result, the crack delay effect of the reinforcing element 103 is minimal.